DTIC ADA461808: F/A-18C to E Wing Morphing Study for the Abrupt Wing Stall Program

31 Pages·1.4 MB·English

by  Defense Technical Information Center

AIAA 2003-0925

F/A-18C TO E WING MORPHING STUDY
FOR THE ABRUPT WING STALL
PROGRAM

Bradford E. Green
NAVAIR
Patuxent River, MD

James D. Ott
Combustion Research and Flow Technology, Inc.
Dublin, PA

41st AIAA Aerospace Sciences Meeting & Exhibit
6-9 January 2003
Reno, Nevada

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F/A-18C to E Wing Morphing Study for the Abrupt Wing Stall Program
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Standard Form 298 (Rev. 8-98)
Prescribed by ANSI Std Z39-18
F/A-18C TO E WING MORPHING STUDY
FOR THE ABRUPT WING STALL PROGRAM

Bradford E. Green *
NAVAIR
Patuxent River, MD

James D. Ott †
Combustion Research and Flow Technology, Inc. (CRAFT)
Dublin, PA

Abstract Nomenclature

In an effort to determine the impact of various a free-stream speed of sound, ft/s
∞
wing parameters on the abrupt wing stall b wing span, ft
phenomenon encountered by the pre-production c chord of wing airfoil section, ft
F/A-18E, various characteristics of the F/A-18C C aircraft drag coefficient, D/(q S)
D ∞
wing were modified to reflect the design changes C aircraft lift coefficient, L/(q S)
L ∞
incorporated into the F/A-18E wing. The
C static pressure coefficient, (P-P )/(q )
p ∞ ∞
parameters evaluated during this study included
C wing-root bending-moment coefficient
WB
thickness, camber, twist, leading-edge radius,
c sectional lift coefficient, l/(q c)
l ∞
leading-edge flap-chord ratio and the addition of a
D drag on aircraft, lb
leading-edge snag. The wing parameters were
L lift on aircraft, lb
modified independently and then in combination
l wing sectional lift, lb/ft
to determine their impact on the abrupt stall. M free-stream Mach number, U / a
∞ ∞
Several potential computational Figures of Merit
P static pressure, lb/ft2
were evaluated to determine their utility for the q dynamic pressure, (0.5ρU 2), lb/ft2
∞ ∞ ∞
prediction of an abrupt wing stall. One of the
Re Reynolds number, (ρU c )/µ
most promising Figures of Merit for indicating the ∞ ∞mac ∞
S wing reference area, ft2
onset of an abrupt stall was found to be the wing-
U free-stream velocity, ft/s
root bending-moment coefficient. Using this ∞
x chordwise distance along wing, ft
Figure of Merit, it was determined that the
y spanwise distance along wing, ft
incorporation of a leading-edge snag, the

reduction of leading-edge flap-chord ratio and the
Greek
elimination of camber are the likely contributors

to the abrupt stall phenomenon encountered by the
α angle of attack, degrees
aircraft.
ρ∞ free-stream density, slugs/ft3
µ∞ free-stream dynamic viscosity, slugs/ft-s

Subscripts

* Aerospace Engineer, Member AIAA mac mean aerodynamic chord
† Aerospace Engineer, Member AIAA ref reference
This paper is declared a work of the U.S. root wing root
Government and is not subject to copyright tip wing tip
protection in the United States. ∞ free-stream
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American Institute of Aeronautics and Astronautics
production configuration without the wing
Acronyms modifications. In particular, the current study was
for an aircraft without porosity.
AEDC Arnold Engineering Development There are significant differences in the wing
Center characteristics between the F/A-18C and the F/A-
AoA Angle of Attack 18E. During the development of the F/A-18E,
AWS Abrupt Wing Stall changes were made to the wing by modifying the
CFD Computational Fluid Dynamics thickness, camber, twist, leading-edge radius,
EMD Engineering and Manufacturing leading-edge flap-chord ratio and adding a
Development leading-edge snag. The differences in these wing
FOM Figure of Merit parameters between the F/A-18C and the F/A-18E
LEX Leading-Edge Extension are shown in Table 1. Since the F/A-18C does not
MAC Mean Aerodynamic Chord experience abrupt stall within its flight envelope,
PWT Propulsion Wind Tunnel the goal of this research was to determine which of
WRBM Wing-Root Bending-Moment these wing parameters contributed to the abrupt
stall phenomenon initially encountered by the F/A-
18E.
Introduction Computational Fluid Dynamics (CFD) was
used as the tool to accomplish this goal. To
During the Engineering and Manufacturing determine the effect of each of the wing
Development (EMD) phase of the F/A-18E/F parameters listed in Table 1, the grid of the F/A-
program, pre-production aircraft encountered an 18C was modified or “morphed” to reflect the
uncommanded, abrupt rolling motion during changes made to these parameters when designing
transonic maneuvering conditions. Commonly the F/A-18E. For example, to determine the effect
referred to as abrupt wing stall (AWS), this of twist on abrupt stall, the twist was removed
phenomenon occurs when there is a loss in roll from the F/A-18C grid to form a “morphed” grid.
damping or when the amount of flow separation The twist was removed from the F/A-18C grid
on the upper surface of the wings increases rapidly since, as shown in Table 1, the F/A-18C wing has
and asymmetrically. A rolling moment results twist while the F/A-18E wing does not. The
from the difference in lift that is generated parameters listed in Table 1 were modified
between the two wings. Abrupt stall degrades the independently and then in combination.
tracking capability of the aircraft and causes safety
concerns if it occurs at low altitudes.1 Table 1. Comparison of wing parameters
Modifications of the automatic leading-edge on the F/A-18C and F/A-18E
flap schedule and the addition of a porous surface Wing
F/A-18C F/A-18E
at the wing fold solved the abrupt stall problem on Parameter
the F/A-18E/F and were incorporated into the
3.5% to 3.8% to
Thickness
production versions of the aircraft. As a result of
5.0% chord 6.2% chord
the modifications, the F/A-18E/F’s being delivered
to the fleet are no longer susceptible to AWS. On Camber 0.6% chord No Camber
the recommendation of a high level panel that over
viewed the technical problems encountered in the Twist 4° Outboard No Twist
early development program, the AWS program Twist
was formed to obtain a better understanding of the
Snag No Snag LE Snag
causes of abrupt stall.2 The AWS program is a
national team composed of the Navy, NASA, Air 0.08 to 0.31
LE radius 0.04 inches
Force and several universities. The research
inches
presented in this paper was funded through the
AWS program. LE flap-chord 20% chord 11.3% to
ratio 18.1% chord
The reader is cautioned that references to the
F/A-18E configuration herein are to the pre-
2
American Institute of Aeronautics and Astronautics
After analyzing each of the configurations at with wind-tunnel results than did the results using
transonic speeds, the wing-root bending-moment Spalart-Allmaras.2
(WRBM) coefficient was plotted as a function of
angle of attack (AoA). Based on the Figures of
Merit developed during the AWS program3, abrupt F/A-18C Grid
stall can occur when the slope of the coefficient of
WRBM curve changes sign. The AoA at which The structured, overset F/A-18C grid is shown
the slope changes sign is of particular importance. in Figure 1. The grid is comprised of 79 blocks
If, for any of the morphed configurations, the and has approximately 13.9 million points. To
slope of the WRBM curve changes sign at a lower reduce the size of the grid, only half of the
AoA than it does for the F/A-18C, then this configuration is modeled using a plane of
indicates that the particular wing parameter or symmetry along the fuselage centerline. The
parameters being modified may be contributing to horizontal and vertical tails are not represented in
abrupt stall. the grid, since it was anticipated that they would
In the following sections, the CFD flow solver have little impact on the abrupt stall phenomenon
and the F/A-18C grid will be discussed. Next, the encountered by the F/A-18E.
F/A-18C results from CFD will be compared to In the grid, the leading- and trailing-edge flaps
the results from the wind-tunnel experiment. are deflected to 6 and 8 degrees, respectively,
Then, the WRBM Figure of Merit will be while the aileron is not deflected. This is referred
discussed as well as the methodology used to to as a 6°/8°/0° flap setting. A missile launcher
accomplish the goal of this research. The CFD and a Sidewinder missile are attached to the wing
analysis of the F/A-18C grid will be presented tip.
prior to discussing the results of the morphing The WIND flow solver required 900 MBytes
study. In the last section, the concluding remarks of memory on Silicon Graphics Octanes and
will be presented. Origin 2000/3000/3800 machines. Each solution
required approximately 6000 CPU-hours to
complete.
Discussion of the CFD Flow Solver

The WIND flow solver4 was used to obtain the Comparison with Wind-Tunnel
flow solution on the F/A-18C configuration and all Experiment
of the morphed configurations studied during this

project. This flow solver is an extension of the
The F/A-18C grid was analyzed at Mach 0.8
NASTD code, which was developed by
and 0.9 at a Reynolds number of 2.07 million
McDonnell Douglas Aerospace-East. WIND
based on the mean aerodynamic chord (MAC) for
solves the Reynolds-averaged Navier-Stokes
several AoA and the results were compared to the
equations in conservative form using second-
wind-tunnel data obtained from the 4-Foot
order-accurate finite differences on structured,
Transonic Propulsion Wind Tunnel (PWT) at the
multi-zonal computational grids. The explicit
Arnold Engineering Development Center (AEDC).
terms are computed using either upwind or central
A Reynolds number of 2.07 million corresponds to
difference. The implicit terms are computed using
the wind-tunnel Reynolds number of 3 million/ft
either an approximately factored or four-stage
for a 6% scale model. In Figure 2, the lift and
Runge-Kutta scheme.
drag coefficients from CFD are compared to the
Several turbulence models are available within
wind-tunnel data obtained at the AEDC. In
WIND. Some of these models include Cebeci-
general, the lift and drag coefficients from CFD
Smith, Baldwin-Lomax, Spalart-Allmaras and
compare very well with the wind tunnel data. The
Menter’s SST model. Menter’s SST model5 was
lift from CFD does slightly over-predict the wind
used during this research, since it was determined
tunnel data. This over-prediction can probably be
early in the AWS program that the solutions on the
explained by the fact that the wind-tunnel model
F/A-18E using Menter’s SST correlated better
had horizontal and vertical tails, while the CFD
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grid did not. The wind tunnel data shown in lateral activity being present. The change in sign
Figure 2 was obtained while the horizontal tail suggests that further testing is necessary to
deflection angle was 0 degrees. determine if there is an unacceptable lateral
Oil flow images on the wing upper surface problem.
were also obtained during the AEDC wind tunnel Since the change in sign of the slope of the
test. In general, the surface-restricted particle WRBM curve can indicate the onset of the abrupt
traces from CFD compare well with the oil flow stall, the AoA where the slope changes sign will
images from the wind tunnel experiment. Figure 3 be referred to as the onset AoA. The onset AoA
shows two comparisons between wind tunnel oil indicates the lower boundary of an AWS region of
flow images and surface-restricted particle traces interest, in which abrupt stall could occur. The
from CFD at Mach 0.8. At an AoA of 7 degrees, AWS region of interest extends approximately
both the wind tunnel oil flow and CFD particle from the onset AoA to an AoA where the slope of
traces predict flow separation and reattachment on the WRBM starts increasing again.
the wing box. At an AoA of 11 degrees, both the In Figure 4, the WRBM coefficient is plotted
oil flow and CFD surface-restricted particle traces as a function of AoA for the F/A-18C at Mach 0.8
indicate that flow separation has reached the and 0.9 and the F/A-18E at Mach 0.9. It can be
leading edge. seen in the figure that the onset AoA for the F/A-
The correlation of the lift, drag and oil flow 18C is 10 degrees at Mach 0.8 and 11 degrees at
images between CFD and the wind tunnel Mach 0.9. Although the slope of the WRBM
experiment indicate that the grid and flow solver curve changes sign for the F/A-18C with 6°/8°/0°
are adequate for accurately predicting the flow flaps, the flaps are not on schedule at 10 and 11
over the F/A-18C. degrees AoA. As a result, the F/A-18C does not
experience abrupt stall within its flight envelope.
The pre-production F/A-18E, however, did
Wing-Root Bending-Moment Coefficient experience abrupt stall. The WRBM coefficient in
as a Figure of Merit Figure 4 for the F/A-18E with 6°/8°/4° flaps
changes sign at 7.5 degrees. Since the 6°/8°/4°
The WRBM coefficient was utilized as a flaps are nearly on schedule for the pre-production
Figure of Merit (FOM) for identifying AWS.3 The aircraft at 7.5 degrees AoA, the abrupt stall
WRBM is calculated using the equation occurred within the flight envelope of the aircraft.
When analyzing the WRBM coefficient to
y determine if the slope changes sign, smaller
tip
WRBM = ∫l(y− y )dy increments in AoA may be necessary. In Figure 5,
ref
the initial and refined WRBM curves are shown
y
root
for two cases. For the first case, which is
on the wing of the aircraft. In this equation, y illustrated using the plots at the top, the initial
ref
was taken to be at the interface between the wing WRBM curve was obtained using AoA increments
and the leading-edge extension (LEX) of the F/A- of 1 degree. Using increments of 1 degree does
18C. The WRBM coefficient is then obtained not clearly resolve a change in sign of the slope. It
using the equation appears, though, that a change in sign of the slope
could occur between 10 and 11 degrees AoA. As
WRBM a result, the 10.5 degree AoA case was analyzed
C = and used to create the refined WRBM curve
WB q S b
∞ ref ref shown on the top, right-hand side of the figure.
This plot clearly shows that the slope changes sign
Based on correlation between wind tunnel and at an AoA of 10 degrees.
flight tests, abrupt stall can occur when the slope The bottom two plots in Figure 5 are used to
of the WRBM versus AoA curve changes sign. illustrate the second case where a smaller AoA
Abrupt stall, however, is not guaranteed to occur increment could be helpful. The initial WRBM
when the slope of the WRBM changes sign. The curve in the lower, left-hand plot was generated
change in sign could also occur with no significant using AoA increments of 1 degree. The slope of
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American Institute of Aeronautics and Astronautics
this curve changes sign at 7 degrees AoA. When Morphing cases 7 through 11 in Figure 7
the curve is refined by calculating the flow at 7.5 represent cases where two wing parameters were
degrees AoA, the slope changes sign at 7.5 modified simultaneously. Specifically, these cases
degrees, as shown in the plot on the lower, right- show the effect of the latter five wing parameters
hand side of Figure 5. In addition, the refined on a wing with a snag. For example, morphing
curve has an extremely different character since case #8 represents a configuration where a snag
the peak of the curve where the slope changes sign was added to the F/A-18C wing while at the same
is significantly larger. time the camber was removed from the wing. The
The lift-curve slope was also considered as a camber was removed from the wing since the F/A-
potential FOM during the AWS program.3 18E wing is uncambered.
Specifically, the location of the break in the lift After creating the new grids with the desired
versus AoA curve was viewed as a possible changes in wing parameters, solutions were
indicator of abrupt stall. In Figure 6, the lift and generated at Mach 0.8 and 0.9 for several AoA’s.
WRBM coefficients are plotted as a function of The WRBM coefficient was then plotted as a
AoA for the F/A-18C at Mach 0.8 and 0.9. While function of AoA to determine which wing
the slope of the WRBM coefficient changes sign at parameters were having an effect on abrupt stall.
10 and 11 degrees for Mach 0.8 and 0.9, From Figure 4, one can see the changes in
respectively, the slope of the lift coefficient WRBM coefficient that indicate that a particular
changes sign only at 10 degrees for Mach 0.8. At wing parameter is contributing to abrupt stall.
Mach 0.9, it is difficult to determine where the lift Since the F/A-18C wing is being morphed to
versus AoA curve breaks. As a result, the change mimic an F/A-18E wing, one should expect that
in sign of the slope of the WRBM curve was used wing parameters contributing to abrupt stall would
as the FOM for this study. move the onset AoA in the WRBM curve from the
F/A-18C toward the F/A-18E. In particular, the
AoA at which the slope changes sign will be
Methodology reduced from 10 and 11 degrees at Mach 0.8 and
Mach 0.9 to approximately 7.5 degrees AoA or
In an effort to determine their impact on less.
abrupt stall, the wing parameters listed in Table 1 To mimic the flap settings for the F/A-18C
were evaluated by modifying the F/A-18C wing to grid that was analyzed and compared to wind-
reflect the changes that were made to the wing tunnel data, each of the morphing configurations
when designing the F/A-18E. This process of were analyzed with 6°/8°/0° flaps. In Figure 4 and
modifying the wing parameters on the F/A-18C throughout this paper, the F/A-18C with 6°/8°/0°
wing is referred to as “morphing”. Eleven different flaps was compared to the F/A-18E with 6°/8°/4°
morphing configurations, which are shown in flaps. The only difference between these two flap
Figure 7, were used to determine if the six wing settings is that the aileron is not deflected on the
parameters were contributing to the abrupt stall. F/A-18C, while the aileron is deflected –4 degrees
For the first six morphing cases shown in Figure 7, on the F/A-18E. This difference in flap settings is
the F/A-18C was modified to reflect the change probably not contributing to the significant
made for the specific wing parameter. For difference in the AWS characteristics seen
example, morphing case #4 in Figure 7 represents between the two aircraft, since unpublished
the F/A-18C with an untwisted wing. While Veridian wind-tunnel data indicates that abrupt
removing the twist from the wing, the other wing wing stall is more sensitive to the leading-edge
parameters shown in Table 1 were unchanged. flap setting than it is to the trailing-edge flap and
The twist was removed from the F/A-18C wing aileron settings.
since the F/A-18E wing is untwisted, as shown in Before presenting the results of the morphing
Table 1. study, the results of the analysis of the F/A-18C
will be presented.

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Analysis of the F/A-18C configurations and the baseline F/A-18C for Mach
0.8 and 0.9 is shown in Table 2. As an example,
In this section, the baseline F/A-18C results the first case listed in the table represents the F/A-
will be discussed. In Figures 8 and 9, the pressure 18C with reduced leading-edge radius. In general,
coefficients, surface-restricted particle traces and as one goes from the top to the bottom of the table,
regions of off-body flow reversal on the upper the difference between the onset AoA of the
surface of the wing are shown for several AoA at morphed configuration and the F/A-18C becomes
Mach 0.8 and 0.9. The areas above the wing larger.
covered in red indicate the regions of flow
reversal. As expected, for both Mach numbers the
flow separation moves toward the leading edge as Table 2. Difference in onset AoA between
the AoA is increased. Since abrupt stall occurs morphing configurations and the F/A-18C
when the amount of flow separation increases
Mach 0.8 Mach 0.9
rapidly with AoA, analyzing the regions of flow
separation and how they change with increasing
LE Radius 0 0
AoA is important.
In Figure 10, the sectional lift coefficient and
its derivative with AoA are plotted as a function of Twist -1 0
spanwise location for the F/A-18C at Mach 0.8
and 0.9. It can be seen from the figure that the Camber -1.5 0
sectional lift increases over the span until 11 and
12 degrees AoA at Mach 0.8 and 0.9, respectively. 11.3% LE Flap-
-1 -1
The discontinuity in the F/A-18C sectional lift Chord Ratio
plots near 2y/b=0.67 is caused by the different
Thickness -2.5 & 0 -1
deflections between the trailing-edge flap and the
aileron. Recall that the WRBM coefficient for the
F/A-18C changed sign at 10 and 11 degrees for Snag -2 -2
Mach 0.8 and 0.9, respectively. This is consistent
with the loss in lift that appears in Figure 10. Snag, Thickness -2 -1
Plots of dc/dα yielded valuable information
l
during this research. These plots give insight into Snag, Twist -2 -2
the location on the span where lift is being lost.
This can help tremendously when determining the
Snag, LE Radius -2 -2
factors that are contributing to abrupt stall.
Furthermore, there is a good correlation between
Snag, Camber -2.5 -2.5
the AoA where the largest amount of lift is lost on
the dcl/dα plots and the onset AoA from the Snag with
-3.5 -3.2
WRBM curves.
Tapered LE Flap

Results
Snag

In this section, the results from the morphing
During the design of the F/A-18E wing, a
study will be presented. In each of the six
leading-edge snag was added to the wing to
subsections that follow, the results from a
improve the high-lift performance of the aircraft at
particular wing parameter will be discussed and
low speeds. There is no leading-edge snag on the
conclusions will be drawn about the effect of the
F/A-18C.
wing parameter on abrupt stall.
During this research, a leading-edge snag was
To help compare the cases, the difference
added to the F/A-18C wing to determine its effect
between the onset AoA for each of the morphing
on abrupt stall. This case corresponds to morphing
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case #1 in Figure 7. The planform of the F/A-18C The sectional lift and dc/dα are plotted as a
l
wing with a snag is compared to that of the F/A- function of spanwise location for the F/A-18C and
18C wing in Figure 11. In this figure, the baseline F/A-18C with a snag at Mach 0.9 in Figure 14.
F/A-18C wing is shown in the upper-left corner, The dc/dα curves for the F/A-18C with a snag
l
while the F/A-18C wing with a snag is shown in indicate that the largest loss in lift occurs between
the lower-left corner. The wings are colored to 9 and 9.5 degrees AoA and is centered near the
distinguish the leading-edge flap, wing box, location of the snag. Furthermore, this loss in lift
trailing-edge flap and aileron. is greater than the F/A-18C experienced between
The snag was placed at the same non- 11 and 12 degrees AoA.
dimensional spanwise location as it is on the F/A- In Figure 15, the pressures, surface-restricted
18E. To create the snag, the leading-edge flap of particle traces and regions of off-body flow
the F/A-18C was extended so that the ratio of the reversal on the upper surface of the wing for the
chord length between the airfoil section outboard F/A-18C and the F/A-18C with a snag are
of the snag and inboard of the snag was identical compared at Mach 0.9 for 8, 9 and 10 degrees
to that of the F/A-18E. Since the snag was placed AoA. In this figure, the F/A-18C wing is shown
near the wing-fold fairing on the F/A-18C, the on the top while the F/A-18C wing with a snag is
wing-fold fairing and other protuberances were shown on the bottom. At 8 and 9 degrees AoA,
removed from the wing to simplify the process of the flow separates at nearly the same location on
adding the snag. With the exception of the wing- both wings. However, at 10 degrees AoA the flow
fold fairing, the leading-edge flap on the F/A-18C separation on the wing with a snag has moved
wing was only modified outboard of the snag further forward than the separation on the F/A-
location. 18C. This rapid, forward movement of the flow
The results of WRBM coefficient for the F/A- separation region for the F/A-18C with a snag
18C and the F/A-18C with a snag are shown in between 9 and 10 degrees AoA is causing the
Figure 12. It can be seen in the figure that the slope to change sign in the WRBM in Figure 12
addition of a snag changes the WRBM curves and the loss in lift that occurs in Figure 14.
significantly. At Mach 0.8, the onset AoA for the In Figure 16, the chordwise pressure
configuration with a snag is 8 degrees. This distribution is plotted just inboard of the snag for
represents a reduction of two degrees in the onset several AoA for the F/A-18C and F/A-18C with a
AoA from the F/A-18C. At Mach 0.9, the addition snag at Mach 0.9. For the F/A-18C with a snag,
of the leading-edge snag reduces the onset AoA by the pressure coefficient is more negative at the
two degrees from 11 to 9 degrees. A 2-degree leading edge and the pressure gradient on the
reduction in the onset AoA is large and, as a result, leading-edge flap is more adverse than it is on the
the snag is most likely a significant contributor to F/A-18C. Between 9 and 12 degrees AoA, the
the abrupt stall initially encountered by the F/A- shock on the upper surface of the F/A-18C with a
18E. snag moves forward more rapidly than it does on
The WRBM curves at Mach 0.8 and 0.9 for the F/A-18C. At 12 degrees AoA, the flow
the F/A-18C with a snag are compared in Figure separation has reached the leading edge of the
13. These curves differ in character, partially wing with a snag.
because the WRBM curve at Mach 0.8 has a As mentioned above, the leading-edge snag is
sharper peak where the slope changes sign. The contributing significantly to abrupt stall at both
sharper peak at Mach 0.8 is also manifested by a flow conditions considered. The reason that the
larger loss in lift in the dc/dα curves, which are snag is contributing to abrupt stall is shown in
l
also shown in Figure 13. In each of the dc/dα Figures 17 and 18. The particle traces near the
l
plots, the largest loss in lift corresponds to the snag on the F/A-18C with a snag configuration are
onset AoA on the WRBM plots. These results shown in Figure 17. As the flow accelerates
indicate that the dc/dα curves are useful for around the inboard edge of the snag, a region of
l
predicting the severity of the abrupt stall, in low pressure is formed on the upper surface of the
addition to the AoA where the abrupt stall occurs. leading-edge flap inboard of the snag. This region
of low pressure causes a steep adverse pressure
gradient which separates the flow on the flap
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American Institute of Aeronautics and Astronautics
inboard of the snag. The region of reversed flow WRBM changes sign. This increases the
is shown in red on the right-hand plot in Figure 17. credibility that dc/dα can be used as a FOM.
l
As a result of the adverse pressure gradient and For the second case, the F/A-18C was
flap separation, the boundary layer thickens and modified by adding a leading-edge snag and
separates earlier on the wing. A typical plot of the thickening the wing. The WRBM for this
contours of boundary layer thickness for the F/A- configuration is compared to that of the F/A-18C
18C and F/A-18C with a snag is shown in Figure with a snag at Mach 0.8 and 0.9 in Figure 21. At
18. In this figure, the boundary layer becomes Mach 0.8, the increase in thickness does not
thicker as the color changes from blue to red. For change the onset AoA. At Mach 0.9, however, the
the F/A-18C with a snag, the boundary layer starts onset AoA increases by 1 degree as a result of the
thickening on the leading-edge flap just inboard of increase in wing thickness. This was the only case
the snag. examined during this study where a change in a
Since the addition of a leading-edge snag to wing parameter increased the onset AoA.
the F/A-18C wing reduced the onset AoA by 2 Based on these results, thickness does not
degrees at both Mach 0.8 and 0.9, the camber, appear to be contributing to abrupt stall on the
thickness, twist, leading-edge radius and leading- F/A-18E.
edge flap-chord ratio on the F/A-18C were
modified independently and then in combination Camber
with the leading-edge snag. The results for these
five wing parameters in combination with the snag Morphing cases 3 and 8 in Figure 7 were used
are compared to the results of the F/A-18C with to analyze the effect of camber on abrupt stall. In
the snag in this section. the first case, the camber was removed from the
F/A-18C wing. In the second case, the camber
Thickness was removed from the F/A-18C wing while also
adding a leading-edge snag. The WRBM curves
The effect of thickness on abrupt stall was for these cases at Mach 0.8 and 0.9 are shown in
determined by analyzing two different Figure 22. The results for the first case are shown
configurations, as shown in Figure 7. In the first in the plots at the top, where the WRBM curves
case, the thickness of the F/A-18C wing was for the F/A-18C with an uncambered wing are
increased, as indicated in Table 1. In Figure 19, compared to that of the F/A-18C. These results
the WRBM coefficient is shown as a function of indicate that the removal of the camber from the
AoA for the F/A-18C and the F/A-18C with a F/A-18C wing reduced the onset AoA by 1.5
thickened wing for Mach 0.8 and 0.9. At Mach degrees at Mach 0.8. The removal of the camber
0.8, the slope changes sign at 7.5 and 10 degrees from the wing does not change the onset AoA at
for the thickened wing. The change in sign at 7.5 Mach 0.9.
degrees represents a 2.5-degree reduction in onset The results for the second case are shown in
AoA from the F/A-18C. At Mach 0.9, the slope the plots at the bottom of Figure 22. In these
changed sign at 10 degrees for the thickened wing. plots, the WRBM curves for the F/A-18C with a
This represents a 1-degree reduction from the F/A- snag and uncambered wing are compared to those
18C. of the F/A-18C with a snag for Mach 0.8 and 0.9.
The F/A-18C with a thickened wing at Mach These two plots indicate that removal of the
0.8 was the only configuration in this study where camber from the wing with a snag reduces the
the slope changed sign in two distinct AoA onset AoA by 0.5 degree at both Mach numbers.
regions. To test a potential dc/dα FOM, the Based on these results, the removal of the
l
WRBM curve and dc/dα for the thickened wing at camber from the wing could be a contributing
l
Mach 0.8 were plotted as a function of spanwise factor to AWS. A possible explanation for the
location in Figure 20. This figure indicates that contribution of the camber to abrupt stall can be
there is a large loss in lift at 7.5 and again at 10 seen in Figure 23. In this figure, the upper surface
degrees AoA, the two AoA where the slope of the pressure contours, surface-restricted particle traces
and regions of flow reversal for the F/A-18C with
a snag and the F/A-18C with a snag and
8
American Institute of Aeronautics and Astronautics